1. Field of the Invention
This invention relates to thrusters intended to be used for low-power applications such as in the orbital positioning of spacecraft. The thruster disclosed is of the type using a propellant such as hydrazine in which the propellant is heated to a desired temperature prior to exiting through a rocket propulsion nozzle. The heating provides a high specific impulse and facilitates decomposition and/or combustion.
These thrusters are normally used during the lifetime of a three axis stabilized or spin stabilized satellite (presently 8-10 years for synchronous orbit satellites) in order to place in, to change or to maintain orbit station. Satellite on-board propulsion is frequently required to finalize and, in some instances, make major corrections to achieve final orbit circularization and/or orbit station. When this is accomplished with a typical hydrazine-fueled engine, large quantities of propellant may be expanded. Use of a performance augmented engine (using electric energy to extend the nominal chemical reaction performance level) for this function would conserve and retain more fuel for on-orbit functions. Typically, excess electric power is available on a spacecraft even during orbit/station insertion maneuvers. Correction firings are time-spaced, with off periods between firings so as to permit battery recharge for subsequent firings. By this augmentation process, fuel usage can be reduced by as much as 30 percent.
The thruster may also be used for correcting a satellite orbit which has decayed, or for repositioning the satellite to another location or station. Such thrusters can also be used for propelling satellites which follow other satellites or for evading tracking satellites. Another application of the performance-enhanced engine is to change the orbital path of a satellite in order to evade ground tracking or to make ground tracking more difficult. An application of this would be in satellite maneuvering for the sole purpose of decoying or saturating would-be tracking capabilities.
In usage, this engine could be ground-controlled by the spacecraft operating agency, or in some instances of covert operation may be preprogrammed for on-orbit automatic control.
2. Brief Description of the Prior Art
Liquid propellant fueled spacecraft engines operate at performance levels limited by the chemical reaction energy of the propellant. Performance is generally maximum for steady state operating periods of more than a minute and reduced for pulsing operation. For a monopropellant fueled engine (i.e. hydrazine) either a catalyst bed or a thermal decomposer is used to initiate the exothermic reaction process. Of these processes the catalyst bed is the most common. The thermal decomposer is typically brought to operating temperature by means of an electrical resistance heater. These decomposers serve only to initiate the chemical reaction, but do not add to or augment the chemical performance level. To extend the performance level, investigators have suggested use of electrical resistance heaters to boost the chemical performance level by exposing the chemically reacted or reacting propellant to a high temperature heater, thereby increasing the propellant temperature prior to expansion of the propellant through a nozzle. The inherent problem of such a device is the direct contact between the heater and the flow requires that the propellant be as free as possible of contaminants (standards in excess of those typical of most rocket engine usage specification levels). This also precludes use of the more complex propellant formulations such as any that would contain carbon or oxygen, due to possible chemical interactions with the heating element.
In the prior art, the heaters were designed for maximum output during propellant contact. Without heat removal by the propellant the heater would attain excessive temperature and heater burn-out could occur. Accordingly, the power could only be switched on when propellant was flowing, and this meant that the successful transfer of energy from the heating filament to the propellant could be accomplished only when a sufficient amount of electric power was available for heating the propellant at the rate that the propellant was being utilized for thrust. This not only prevented operation during times when battery power was substantially low, but also precluded prehating the thruster with the heater before propellant flow was initiated. This also places limitations upon attainable temperatures. Such an engine cannot be off-flow modulated or pulsed with off period greater than a few milliseconds and as such is limited in its usefulness.
A further characteristic of flow-coupled devices is that the heater is subjected to all pressure fluctuations of the propellant supply and reaction process. Gas dynamic forces from any propellant reaction instability will be transmitted to the heater and may cause a heater distortion failure. Since the flow is circulating through the heater region there is constant flow impacting and washing the heater.
This prior art design also requires use of high temperature sealed electrical feed-through(s) into the chamber. This places restrictions on the overall engine design as to operating temperature and power.
The prior art thrusters used an outer shield having a low emissivity surface in order to reflect as much heat as possible back to within the heating portion of the thruster. This minimized energy losses by maintaining a higher temperature within the heating section of the thruster. Because the minimization of heat loss was accomplished in part by minimizing exterior surface cooling, the exterior tended to remain hot causing heat to be transferred through the thruster's supporting structure to the satellite proper. A further disadvantage of having a high exterior temperature was that infrared sensors could easily distinguish a warm satellite's components from the surrounding space. The rocket nozzle section of thrusters also presented a source of high temperature emissions. This resulted from the high temperatures present at the nozzle's throat and internal expansion chamber areas, which high temperatures were conducted as heat to the outer portions of the rocket nozzle.
In prior augmentation designs, the liquid propellant is first decomposed, vaporized and reacted in a separate chamber allowing some of the reaction energy to be lost.